1. Field of the Invention
The present invention relates generally to gas turbine engine, and more specifically to turbine rotor blade with blade tip cooling.
2. Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98
A gas turbine engine, such as a large frame heavy duty industrial gas turbine (IGT) engine, includes a turbine with one or more rows of stator vanes and rotor blades that react with a hot gas stream from a combustor to produce mechanical work. The stator vanes guide the hot gas stream into the adjacent and downstream row of rotor blades. The first stage vanes and blades are exposed to the highest gas stream temperatures and therefore require the most amount of cooling.
The efficiency of the engine can be increased by using a higher turbine inlet temperature. However, increasing the temperature requires better cooling of the airfoils or improved materials that can withstand these higher temperatures. Turbine airfoils (vanes and blades) are cooled using a combination of convection and impingement cooling within the airfoils and film cooling on the external airfoil surfaces.
The turbine rotor blades have blade tips that form a gap with a blade outer air seal (BOAS) on the stationary housing. This blade tip gap varies in spacing due to engine operation. Hot gas flow will leak through the gap and cause erosion damage to the tip that eventually wears away pieces of the tip that will then further increase the tip leakage flow, which then further causes additional erosion damage.
Prior art blade tips are cooled by drilling holes into the upper extremes of a serpentine flow cooling circuit formed within the airfoil of the blade with cooling holes that open onto the pressure and suction side surfaces just below the blade tip corners along the blade tip edge and on top of the blade tip floor that opens into a squealer pocket. As a result of this cooling design, cooling flow distribution and pressure ratios across these film cooling holes for the airfoil pressure and suction sides as well as the tip cooling holes are predetermined by the internal cavity pressure. In addition, the blade tip region is subject to severe secondary flow field which therefore requires a large number of film cooling holes and cooling flow required for the cooling of the blade tip periphery.